1. Field of the Invention
The present invention relates generally to fluid reaction surfaces, and more specifically to a turbine blade with film cooling slots.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
Rotor blades in a turbine of a gas turbine engine are cooled by passing cooling air through an internal cooling circuit with film cooling holes on the external surface of the airfoil to provide film cooling to the surface. A film cooling hole will open into a diffuser in order to slow the flow such that the film of cooling air develops on the airfoil surface. The cooling film provides a blanket-like effect to keep the hot gas flow from contacting the airfoil surface. The angle which the axis of the film cooling hole makes with the airfoil surface and its relation to the direction of hot gas flow over the airfoil surface at the hole exit are important design factors which influence film cooling effectiveness. The film cooling effectiveness decreases rapidly with the distance from the cooling hole exit. Maintaining a high film cooling effectiveness for as long a distance from the exit hole as possible over as large a surface area as possible is the main goal of airfoil film cooling.
It is well known in the art that engine airfoils must be cooled using a minimum amount of cooling air, since the cooling air is working fluid which has been extracted from the compressor and is therefore unavailable to perform useful work in the turbine. This bleed off from the compressor reduces the engine efficiency. Thus, it is a design challenge to provide the maximum amount of cooling with the minimum amount of cooling air.
U.S. Pat. No. 3,527,543 issued to Howald on Sep. 8, 1970 entitled COOLING OF STRUCTURAL MEMBERS PARTICULARLY FOR GAS TURBINE ENGINES shows a turbine blade with divergently tapered cooling passages of circular cross section to increase the entrainment of coolant in the boundary layer from a given passage. The passages are preferably oriented in a plane extending in the longitudinal direction or partially toward the gas flow direction to spread the coolant longitudinally upon its exit from the passage as it moves downstream.
The velocity of the air leaving the cooling passage is dependent on the ratio of its pressure at the passage inlet to the pressure of the gas stream at the passage outlet. In general, the higher the pressure ratio the higher the exit velocity. Too high an exit velocity results in the cooling air penetrating into the gas stream and being carried away without providing effective film cooling. Too low a pressure ratio will result in gas stream ingestion into the cooling passage causing a complete loss of local airfoil cooling. Total loss of airfoil cooling usually has disastrous results, and because of this a margin of safety is usually maintained. This extra pressure for the safety margin drives the design toward the high pressure ratios. Tolerance of high pressure ratios is a desirable feature of film cooling designs. Diffusion of the cooling air flow by tapering the passage, as in the Howald patent discussed above is beneficial in providing this tolerance, but the narrow diffusion angles taught therein (12 degree maximum included angle) require long passages and, therefore, thick airfoil walls to obtain the reductions in exit velocities often deemed most desirable to reduce the sensitivity of the film cooling design to pressure ratio. The same limitation exists with respect to the trapezoidally shaped diffusion passages described in U.S. Pat. No. 4,197,443 issued to Sidenstick on Apr. 8, 1908 entitled METHOD AND APPARATUS FOR FORMING DIFFUSED COOLING HOLES IN AN AIRFOIL. The maximum included diffusion angles taught therein in two mutually perpendicular planes are 7 degree and 14 degree, respectively, in order to assure that separation of the cooling fluid from the tapered walls does not occur and the cooling fluid entirely fills the passage as it exits into the hot gas stream. With such limits on the diffusing angles, only thicker airfoil walls and angling of the passages in the airfoil spanwise direction can produce wider passage outlets and smaller gaps between passages in the longitudinal direction. Wide diffusion angles would be preferred instead, but cannot be achieved using prior art teachings.
U.S. Pat. No. 4,653,983 issued to Vehr on Mar. 31, 1987 entitled CROSS-FLOW FILM COOLING PASSAGES shows a film cooling hole having a metering hole leading into a diffuser, where the diffusing portion includes a pair of adjoining surfaces which are both parallel to a central axis of the metering hole and another pair of adjoining surfaces which diverge from the central axis, the diverging pair of surfaces being located on the downstream side of the passage in order to provide an improved film cooling flow.
Axial shaped diffusion film cooling holes are normally used for the cooling of a turbine blade suction wall. The use of axial oriented film cooling holes on the suction surface is primarily for the injection of cooling air to be inline with the main stream flow of hot gas over the airfoil surface which is accelerated in the axial direction. Particles such as sand that enter the engine pass through the combustor and heated to the point of becoming a hot liquid. These hot liquid particles of sand then pass into the turbine. However, at the airfoil suction surface downstream of the leading edge region, hot and heavy particles are traveling at the combination of wheel speed (same as the turbine rotation) and also moving in the axial direction. Due to centrifugal loading, the resultant direction of travel of these particles is in the combination of radial and axial directions as shown in FIG. 1. some of the hot and heavy particles at the lower blade span will travel radially outward at a certain angle relative to the blade suction surface depending on where the hot and heavy particle is in relation to the blade span height. In this particular region, called the Impact Zone and represented by reference numeral 14 in FIG. 1, the hot and heavy particles will hit the airfoil suction surface substantially normal to the airfoil surface (represented by arrow V in FIG. 1) and solidify onto the relatively cold airfoil wall. If an axial film cooling hole is used in the impact zone 14, the particles will strike onto the airfoil surface in between the film cooling holes. With the increasing occurrence of particle strikes, the accumulated particles will plug the film cooling hole and block the flow, resulting in no film cooling on the airfoil surface from that hole.
It is therefore an object of the present invention to provide a film cooling hole arrangement on the suction side of a rotor blade that will reduce the chance of film cooling hole plugging in the impact region of the blade.
It is another object of the present invention to provide for a more effective film cooling effect on the suction side of the rotor blade in the impact region.